1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine rotor blade with a serpentine flow cooling circuit.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
Turbine airfoils, such as rotor blades and stator vanes, pass cooling air through complex cooling circuits within the airfoil to provide cooling from the extreme heat loads on the airfoil. A gas turbine engine passes a high temperature gas flow through the turbine to produce power. The engine efficiency can be increased by increasing the temperature of the gas flow entering the turbine. Therefore, an increase in the airfoil cooling can result in an increase in engine efficiency.
Prior art airfoil cooling of blades makes use of a single five-pass aft flowing serpentine cooling circuit. One such prior art 5-pass serpentine flow circuit for an airfoil 10 is shown in FIGS. 1a and 1b and includes a first up-pass channel 11 of the 5-pass serpentine flow circuit near the airfoil leading edge. A showerhead arrangement of film cooling holes 16 is included in the first up-pass channel 11 of the serpentine flow cooling channel to provide film cooling for the high heat load section of the airfoil nose. The cooling air flows into a first down-pass channel 12 downstream from and adjacent to the first up-pass channel 11, and then into a second up-pass channel 13 and a second down-pass channel 14 before entering a trailing edge up-pass channel 15 where the cooling air is finally discharged through a row of trailing edge cooling holes 17. The five channels 11-15 that form the 5-pass serpentine flow cooling circuit of FIG. 1 each extend from the pressure side wall to the suction side wall such that each channel provides near wall cooling for both sides of the airfoil (the pressure side and the suction side).
In the prior art 5-pass aft flowing serpentine cooling circuit of FIG. 1, the internal cavities are constructed with internal ribs connecting the airfoil pressure and suction walls. In most of the cases, the internal cooling cavities are at low aspect ratio which is subject to high rotational affect on the cooling side heat transfer coefficient. In addition, the low aspect ratio cavity yields a very low internal cooling side convective area ratio to the airfoil hot gas external surface.
The object of the present invention is to provide for a blade with a cooling circuit that provides for a near wall spiral flow cooling arrangement which optimizes the airfoil mass average sectional metal temperature to improve airfoil creep capability for a blade cooling design.
Another object of the present invention is to maximize the airfoil cooling performance for a given amount of cooling air and minimize the Coriolis effects due to rotation on the airfoil internal cavities heat transfer performance.